The disclosure relates to gas turbine engines. More particularly, the disclosure relates to turbine blades.
FIG. 1 schematically illustrates a gas turbine engine 20. The exemplary gas turbine engine 20 is a two-spool turbofan having a centerline (central longitudinal axis) 500, a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section 22 drives air along a bypass flowpath 502 while the compressor section 24 drives air along a core flowpath 504 for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it is to be understood that the concepts described herein are not limited to use with turbofan engines and the teachings can be applied to non-engine components or other types of turbomachines, including three-spool architectures and turbines that do not have a fan section.
The engine 20 includes a first spool 30 and a second spool 32 mounted for rotation about the centerline 500 relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
The first spool 30 includes a first shaft 40 that interconnects a fan 42, a first compressor 44 and a first turbine 46. The first shaft 40 is connected to the fan 42 through a gear assembly of a fan drive gear system (transmission) 48 to drive the fan 42 at a lower speed than the first spool 30. The second spool 32 includes a second shaft 50 that interconnects a second compressor 52 and second turbine 54. The first spool 30 runs at a relatively lower pressure than the second spool 32. It is to be understood that “low pressure” and “high pressure” or variations thereof as used herein are relative terms indicating that the high pressure is greater than the low pressure. A combustor 56 (e.g., an annular combustor) is between the second compressor 52 and the second turbine 54 along the core flowpath. The first shaft 40 and the second shaft 50 are concentric and rotate via bearing systems 38 about the centerline 500.
The core airflow is compressed by the first compressor 44 then the second compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the second turbine 54 and first turbine 46. The first turbine 46 and the second turbine 54 rotationally drive, respectively, the first spool 30 and the second spool 32 in response to the expansion.
U.S. Pat. No. 6,331,217 shows a single-crystal airfoil transient liquid phase (TLP) bonded to a polycrystalline root. U.S. Pat. No. 7,762,309 shows casting of a blade having a single-crystal root and a columnar airfoil.